1. Field of the Invention
The present invention relates to a mobile blade of a gas turbine engine in which bending moments resulting at the blade root during operation are compensated, and is applicable both to the mobile blades of axial compressors or turbines and to propellers.
2. Summary of the Prior Art
Known aerodynamic studies of the definition of the profile of the aerodynamic portion or vane of a mobile blade have led to the provision of a geometric locus of the centers of gravity of successive cross-sections of the blade vane represented by a radial straight line passing through the longitudinal axis of rotation of the engine and contained in the median plane through the root and the platform of the blade. FIG. 1 of the accompanying drawings illustrates this arrangement diagrammatically, showing the locus of the centers of gravity of the vane cross-sections as a straight line 1 meeting the axis A of the engine and centered on the platform 2 and the blade root 3.
However, this arrangement does not meet the mechanical demands of the behavior of such a blade in operation under the action of internal stresses likely to induce deformations and due to aerodynamic forces resulting from gas stresses, to centrifugal forces, and to stresses resulting from the bending and torsion torque effects of the vane profile. The presently known solutions for defining a blade vane profile generally lead to inequality of the load distribution on the two sides of the blade root. This asymmetrical distribution of the stresses transmitted to the disc which carries the blades of a mobile stage, resulting from the forces exerted on the blade vane, does not enable the best advantage to be taken of the mechanical characteristics of disc resistance.
However, improvements have already been proposed. French Specification No. 2 556 409 discloses a blade with low centrifugal stresses in which the geometric locus of the centers of gravity of successive vane cross-sections is non-linear and includes two parts of opposite inclinations relative to a radial straight line. FIG. 2 of the drawings illustrates this solution diagrammatically, the line 1a representing the locus of the centers of gravity relative to the platform 2a and the root 3a of the blade. Another proposed solution illustrated diagrammatically in FIG. 3 exhibits a non-linear curve 1b for the geometric locus of the centers of gravity of the vane cross-sections between the platform 2b of the blade, on the one hand, and the radially outer tip of the blade on the other hand. The curve 1b in this case possesses a variable inclination which corresponds to the application of a continuous law of compensation for the bending moments.
However, these known methods have drawbacks, especially in certain specific applications. The non-linearity of the profiles obtained, and particularly of the leading and trailing edges of the blade, is the cause of additional difficulties of implementation and also makes checks in operation more difficult. These methods also impose, in order to obtain a satisfactory definition of the blade, repetitions which are often very laborious. In addition, particularly in the case of large-size blades such as turbofan blades, these methods have the disadvantage of appreciably spacing the tip of the blade from a radial position on the axis of the blade root, especially in the axial direction. This brings about clearances that are too large between the blade tip and the corresponding casing, particularly in a flow path with a conical outer wall. Similarly, in the case of blades, especially those of a large size, having lateral wings called fins, it becomes difficult to obtain the correct support between adjacent fins during operation, which interferes with their vibration-damping function.